1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine airfoil with a cooling circuits.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as an aero gas turbine or an industrial gas turbine engine, rotor blades and stator vanes in the turbine are cooled by passing compressed air through the airfoils to provide cooling. Air cooled turbine airfoils are required, especially in the first and second stages, in order to allow for an extremely high hot gas flow temperature into the turbine. The higher the hot gas flow temperature entering the turbine, the higher will be the efficiency of the engine.
Since the compressed air used for cooling the turbine airfoils is typically bleed off air from the compressor, it is also desirable to minimize the amount of cooling air used in order to also increase the engine efficiency. Work is performed on the compressed air for cooling by the compressor, and this work is not used to produce power in the engine. Therefore, it is desirable to design an air cooled turbine airfoil that will use a minimal amount of cooling air while providing a maximum amount of cooling to the airfoil in order to increase the efficiency of the engine.
Airfoils of the stator vanes and rotor blades may use similar cooling features, but suitably modified for the different configurations of the vanes and blades, and their different operation since the vanes are stationary, whereas the blades rotate during engine operation and are subject to considerable centrifugal forces. The hollow airfoils of the vanes and blades typically have multiple radially or longitudinally extending cooling channels therein in one or more independent cooling circuits. The channels typically include small ribs or turbulators along the inner surface of the airfoils which trip the cooling air for enhancing heat transfer during the cooling process.
Typical cooling circuits include serpentine circuits wherein the cooling air is channeled successively through the serpentine legs for cooling the different portions of the airfoil prior to discharge therefrom.
The vanes and blades typically include various rows of film cooling holes through the pressure and suction sidewalls thereof which discharge the spent cooling air in corresponding films that provide additional thermal insulation or protection from the hot combustion gases which flow thereover during operation.
Yet another conventional cooling configuration includes separate impingement baffles or inserts disposed inside the nozzle vanes for impingement cooling the inner surface thereof. The baffles include a multitude of small impingement holes which typically direct the cooling air perpendicular to the inner surface of the vane for impingement cooling thereof. The spent impingement cooling air is then discharged from the vane through the various film cooling holes.
Impingement cooling of turbine rotor blades presents the additional problem of centrifugal force as the blades rotate during operation. Accordingly, turbine rotor blades typically do not use separate impingement baffles therein since they are impractical, and presently cannot meet the substantially long life requirements of modern gas turbine engines.
Instead, impingement cooling a turbine rotor blade is typically limited to small regions of the blade such as the leading edge or pressure or suction sidewalls thereof. Impingement cooling is introduced by incorporating a dedicated integral bridge or partition in the airfoil having one or more rows of impingement holes. Turbine rotor blades are typically manufactured by casting, which simultaneously forms the internal cooling circuits and the local impingement cooling channels.
The ability to introduce significant impingement cooling in a turbine rotor blade is a fundamental problem not shared by the nozzle stator vanes. And, impingement cooling results in a significant pressure drop of the cooling air, and therefore requires a corresponding driving pressure between the inside and outside of the airfoils during operation.
Since the pressure distribution of the combustion gases as they flow over the pressure and suction sides of the airfoils varies accordingly, the introduction of impingement cooling in turbine rotor blades must address the different discharge pressure outside the blades relative to a common inlet pressure of the cooling air first received through the blade dovetails in a typical manner.
Some prior art turbine airfoils include film cooling holes connected to a serpentine flow cooling circuit to provide film cooling to the exterior surface of the airfoil. Cooling air discharged out from the airfoil as film cooling air is not used further downstream within the airfoil for cooling. This type of internal serpentine flow and film cooling uses a large amount of cooling air from the compressor. In some situations such as with thin turbine blades, film cooling is not necessary.
Complex cooling circuitry have been proposed in the prior art for stator vanes and rotor blades to provide maximum cooling with a minimal amount of cooling air. Minor changes in the cooling configuration of these components have significant affect on the cooling performance thereof, which significantly affect the efficiency and performance of the gas turbine engine.
U.S. Pat. No. 5,246,340 issued to Winstanley et al on Sep. 21, 1993 and entitled INTERNALLY COOLED AIRFOIL discloses a turbine airfoil in FIG. 5 with a series of impingement cavities extending along the blade, each extending from the pressure side to the suction side wall, and each connected by passages to provide a cooling air flow path through the series of cavities. The disadvantage to the Winstanley invention is that the series of impingement cavities extends along the blade chordwise length in substantially a straight line. The cavities are quite large which results in large impingement areas. Also, multiple impingement holes are used spanning the rib that separates adjacent cavities. In some of the cavities, the impingement hole does not direct the cooling air jet against the hot wall surface of the cavity but directly into the middle of the cavity.
U.S. Pat. No. 6,837,683 B2 issued to Dailey on Jan. 4, 2005 entitled GAS TURBINE ENGINE AIRFOIL discloses a turbine blade or vane with a series of impingement chambers extending along substantially the whole length of the airfoil, where each chamber is connected by to the adjacent chamber by passageways. The chambers in Dailey do not alternate from pressure side to the suction side of the airfoil, but do provide for the long side on either the pressure side or the suction side of which alternates among the adjacent chambers in the series. In the Dailey patent, the impingement cavities are quite large with respect to the blade size, and the rib separating the adjacent cavities includes multiple impingement holes. Like the above Winstanley et al patent, the large cavities result in large impingement areas.
U.S. Pat. No. 7,097,426 B2 issued to Lee et al on Aug. 29, 2006 and entitled CASCADE IMPINGEMENT COOLED AIRFOIL discloses a turbine blade that includes an airfoil having opposite pressure and suction sidewalls joined together at opposite leading and trailing edges and extending longitudinally from root to tip. A plurality of independent cooling circuits is disposed inside the airfoil correspondingly along the pressure and suction sidewalls thereof. Each circuit includes an inlet channel extending through the dovetail. One of the circuits includes multiple longitudinal channels separated by corresponding perforate partitions each including a row of impingement holes for cascade impingement cooling the inner surface of the airfoil. The Lee et al patent uses multiple separate cooling circuits. One impingement circuit is used on the pressure side and a second impingement cooling circuit is used on the suction side. A separate impingement cooling circuit is used for the leading edge showerhead arrangement. The pressure side and suction side impingement cooling circuits are arranged substantially along a straight line from leading edge to trailing edge, resulting in large cavities with large impingements areas. Also, the impingement holes are positioned such that the impingement jet is directed to flow against the hot wall section of the cavity at low angles which results in lower heat transfer coefficient.
In the above cited prior art [patents, the series of impingement cavities result in a short series of impingement cavities with large impingement areas and impingement cooling air jets that are not at the best angle to produce the highest heat transfer coefficient.
Accordingly, it is desired to provide a turbine rotor blade having improved impingement cooling therein. It is another object of the present invention to provide for a greater number of impingement cavities in the series flow through the blade without increasing the length of the blade.